Radio frequency data downlink for a high revisit rate, near earth orbit satellite system

ABSTRACT

A satellite system operates at altitudes between 100 and 350 km relying on vehicles including a self-sustaining ion engine to counteract atmospheric drag to maintain near-constant orbit dynamics. The system operates at altitudes that are substantially lower than traditional satellites, reducing size, weight and cost of the vehicles and their constituent subsystems such as optical imagers, radars, and radio links. The system can include a large number of lower cost, mass, and altitude vehicles, enabling revisit times substantially shorter than previous satellite systems. The vehicles spend their orbit at low altitude, high atmospheric density conditions that have heretofore been virtually impossible to consider for stable orbits. Short revisit times at low altitudes enable near-real time imaging at high resolution and low cost. At such altitudes, the system has no impact on space junk issues of traditional LEO orbits, and is self-cleaning in that space junk or disabled craft will de-orbit.

CROSS REFERENCE TO RELATED APPLICATIONS

This application is a continuation of U.S. patent application Ser. No.16/925,081, filed Jul. 9, 2020, which is a continuation of U.S. patentapplication Ser. No. 15/868,812, filed Jan. 11, 2018 (U.S. Pat. No.10,715,245), entitled “Radio Frequency Data Downlink For A High RevisitRate, Near Earth Orbit Satellite System”. This application also claimspriority to U.S. Provisional Patent Application Ser. No. 62/430,727,filed Dec. 6, 2016, now expired entitled “A Satellite System” and U.S.patent application Ser. No. 15/439,533 filed on Feb. 22, 2017 (U.S. Pat.No. 10,351,267), entitled “A Satellite System”. The entirety of U.S.patent application Ser. No. 15/868,812, U.S. Provisional PatentApplication Ser. No. 62/430,727 and U.S. patent application Ser. No.15/439,533 are incorporated herein by reference.

BACKGROUND

Satellites are used in many aspects of modern life, including earthobservation and reconnaissance, telecommunications, navigation (e.g.,global positions systems, or “GPS”), environmental measurements andmonitoring and many other functions. A key advantage of satellites isthat they remain in orbit due to their high velocity that creates anoutward centripetal force equal to gravity's inward force. Therefore,once in orbit, they stay there typically for years or decades. See, forexample, FIG. 8 , which graphically illustrates a best and worst casecurve for expected lifetime of orbiting vehicles as a function ofaltitude. Depending on the angle of the orbit, a satellite will be ableto observe a large fraction of the earth's surface at some point intime.

A key parameter for satellites used for earth observation is therelationship between altitude, orbital angle, and constellation size. Athigher altitudes, the satellite will be able to observe a largerpercentage of the earth's surface, however the orbital time will belonger and the instrument package required to effectively cover a largerarea at a longer range will be larger and more complex, on the otherhand, a longer orbital time means that the satellite will appear to bein view of a given point on the earth for a longer period and the numberof satellites required to keep all of the earth in view all of the timedecreases. In order for one satellite to cover the entire surface of theearth, sun synchronous polar orbits are frequently used.

Satellite orbital heights are typically categorized in three broadsegments: low earth orbit (LEO), medium earth orbit (MEO) andgeostationary earth orbit (GEO). The general uses and characteristics ofthese orbits are shown in Table I and represent generally accepted usageof the terms LEO, MEO and GEO. Satellites can orbit at any altitudeabove the atmosphere, and the gaps in altitude shown in Table 1, such asbetween LEO and MEO, are also used, if less regularly. It is also commonthat satellites may orbit in eccentric, non-circular orbits, therebypassing through a range of altitudes in a given orbit.

TABLE I Typical characteristics of common orbits. LEO 400- 6.9-7.8 Earthobservation, Random orbits, 3-10 Y 2,000 sensing, ISS, lifetime, spacejunk telecom issue, little radiation constellations MEO 15,000- 3.5 GPS,GLONASS, Highest radiation (Van 20,000 Earth observation Allen Belt),equatorial to polar orbits GEO 42,000 3.1 Sat TV, high BW Can remainabove same telecom, weather spot on Earth, typically satellitesequatorial orbits

For imaging, the power requirements of the digital optical package anddownlink grows roughly in accordance with a square law for the samedelivered image resolution. GEO satellites are far too high for apractical optical observation package. LEO, on the other hand, allowsfor reasonable optical size and power and is protected from spaceradiation. Most earth imaging satellites operate at lower altitudes,roughly at the altitude of the international space station (ISS) (400km)or higher, up to about 2,000 km. At these altitudes, the size and powerrequirements of the imaging package are much lower for the sameresolution relative to a geostationary orbit, the earth's magnetic fieldshields the satellite from most damaging space radiation, and theatmosphere is sufficiently thin that orbital decay is not a majorproblem. However, the satellite will only be in view of a given sectionof the earth's surface for a few minutes, and at lower altitudes, lineof sight communication may only be possible for a minute or less. Thisrequires a large constellation satellites or accepting a lower “revisit”rate for a given point on the Earth's surface.

Altitudes lower than the international space station (ISS) have theadvantage that the imaging package can again, be substantially reducedin size, weight, and power consumption, which in theory allows for muchlower cost satellites.

However, atmospheric drag becomes a major consideration for orbits belowthe orbit of the ISS—even the ISS requires regular “boosting” to keepits orbit from decaying rapidly, and orbital decay issues growexponentially below ISS altitudes. The assignee of the present inventionhas addressed this issue with a novel low drag orbital vehicle andconstellation design described in co-pending application with Ser. No.15/868,794, entitled “System For Producing Remote Sensing Data From NearEarth Orbit,” to Thomas E. Schwartzentruber and Ronald E. Reedy, andco-pending application with Ser. No. 62/616,325, entitled “AtomicOxygen-Resistant, Low Drag Coatings And Materials,” to Timothy Mintonand Thomas E. Schwartzentruber. This enables Near Earth Orbiters, NEOs,a term we use to describe the system and its constituent vehicles (i.e.,a “NEO satellite system”, “NEO vehicle” or a “NEO satellite”) operatingin stable orbits at 100-350 km. Therefore, it is a purpose of thisinvention to describe a satellite system based on orbital vehiclesoperating in stable Earth orbits at altitudes well below traditionalsatellites, specifically between approximately 100 and 300 km.

The availability of a low drag, low cost, high endurance (multi-yearmissions) satellite suitable for altitudes under 300 km allows for theuse of imaging equipment that is low cost and low power yet achievesresolutions that are currently only available from much more expensive,higher altitude satellites and much higher revisit rates because theconstellation can be large for the same cost. There remains, however,the problem of retrieving the high revisit rate, high-resolution datacollected by a large constellation of NEO satellites. Smallerconstellations at higher orbits solve this problem either because thedownlinks are relatively lower data rates per downlink (voice calls), orusing very high bandwidth, but expensive earth stations. Higher orbitsalso imply long revisit times—for a single satellite, the revisit timemay be measured in days. At sub 300 km orbits, the orbital window for adownlink to a given station is too small to tolerate a small number ofearth stations with the power available in a small satellite.Furthermore, in order to fully take advantage of the high revisit ratespossible in a large NEO constellation, the downlink must be sufficientlyrobust to allow for near real time download of data, with latency on theorder of seconds or at most, a few minutes. In the event that a downlinkis not immediately available, the processor can buffer the captured datafor later transmission when a receiving station is within range.

Consequently, the ground station network must in some respects mirrorthe satellite network, with a large number of ground stations to ensurethat a given satellite can be continuously pushing its image data down.Achieving low power consumption for a given bandwidth is also essential,since the small satellite profile required to achieve a low persatellite cost has a correspondingly small surface area available forsolar panels to generate power for the downlink and antennas to providethe antenna gain.

The combination of these factors means that there is a need for a lowcost, low power, high data rate satellite downlink system that meets themission payload requirements for a small orbital vehicle under 300 kmthat still has a camera resolution on the order of 1 meter, coupled witha corresponding robust network of ground stations to match the capacityof the satellites and reduce the delay time in retrieving a highresolution image of a given point on the earth to sub one hour times.

SUMMARY

In one example, a downlink is described that operates in the Ku band andworks with a satellite network of imaging devices with a roughly 1 meterresolution imaging 240 square km per second.

In another example, a downlink is described that operates with an earthnetwork that uses approximately 70 earth stations to cover thecontinental United States and to provide continuous downlink coveragewithin that area.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 shows an example overall design of a low profile satellite usedwith the present invention.

FIG. 2 shows a perspective view of an example low profile satellite usedwith the present invention.

FIG. 3 shows a cross-section of an example low profile satelliteillustrating various components for use with the present invention.

FIG. 4 shows an example antenna design for a low cost earth station foruse with the present invention.

FIG. 5 shows an example earth station spacing plan for use with thepresent invention.

FIG. 6 shows an example of satellites interacting with plural groundstations for use with the present invention.

FIG. 7 shows an example of satellite necklaces for use with the presentinvention.

FIG. 8 graphically illustrates a best and worst case curve for expectedlifetime of orbiting vehicles as a function of altitude.

The several figures provided here describe examples in accordance withaspects of this disclosure. The figures are representative of examples,and are not exhaustive of the possible embodiments or full extent of thecapabilities of the concepts described herein. Where practicable and toenhance clarity, reference numerals are used in the several figures torepresent the same features.

DETAILED DESCRIPTION

This detailed embodiment is exemplary and not intended to restrict theinvention to the details of the description. A person of ordinary skillwill recognize that exemplary numerical values, shapes, altitudes,applications of any parameter or feature are used for the sole purposeof describing the invention and are not intended to be, nor should theybe interpreted to be, limiting or restrictive.

The presented embodiments provide digital, optical imaging systems aswell as other imaging schemes, such as synthetic aperture radar (SAR)and/or thermal imagers. Each system is equally suitable, as long asassociated imaging equipment acquires and generates data at anacquisition and transmission rate comparable with the contemplatedinvention to and from a comparable altitude.

A constellation of NEO imaging orbiters is placed in orbits at analtitude of about 260 km. The vehicles are equipped with digital imagingoptics that will resolve features down to 1 meter in size across a 30 kmwide imaged swath, with an orbital velocity of roughly 8 km/second. Atthis altitude, the useful amount of time that a given satellite isoverhead is about one minute, and if there is no overlap, a single swathwill scan across the equator in about 1350 orbits. The circumference ofthe earth is 40,075 km, so 30 km swaths result in a single satellitebeing in view of a given portion of the earth about once every 1900hours or 80 days at the equator. These times improve towards the poles,so at mid latitudes the revisit rate is substantially higher. Even so,to achieve sub one hour revisit rates for the middle latitudes (NorthAmerica in particular) requires a very large number of satellites, andthe present system is designed to work with a large constellation, from100 to over a 1000 satellites. The constellation is easily scaled, suchthat increasing the number of satellites to 10,000 can be achieved.

A large satellite at 260 km will experience considerable drag, whichrequires a large amount of propellant to keep in orbit, furtherincreasing the size of the vehicle. Accordingly, a 260 km satellite mustbe small and present a low drag profile, such as that disclosed in therelated application entitled “System For Producing Remote Sensing DataFrom Near Earth Orbit.” However, the small size means limited surfacearea for solar panels, which results in lower available power for theradio downlink, which in the present embodiment must operate in nearreal time. Accordingly, the present invention is designed for a powerbudget of 30 watts.

FIG. 1 illustrates an exemplary version of a NEO vehicle 100. Thevehicle is a low drag, “pizza box” design with a wedged leading edge andsolar panels deployed perpendicular to the axis of flight, to minimizethe cross sectional area exposed to forward collisions from atomicoxygen and other particles. The “flat” form factor limits the areaavailable for parabolic or large antenna arrays. Accordingly, theantenna for the downlink radio is constructed in the present inventionas a flat, phased array antenna.

The NEO vehicle 100 can include an electric propulsion engine 106 togenerate thrust by, for example, consuming an ionized fuel to maintainthe desired orbit. Although represented as being external to the vehiclebus 102, the engine 106 can be integrated within the bus 102, shieldedby one or more panels of the bus 100, and/or dimensioned to extendbeyond a surface of the bus 102, in accordance with the presentdisclosure. One or more stabilization surfaces or panels 108 can beemployed, designed to enhance the stability of the NEO vehicle 100, aswell as support solar paneling to collect power. The NEO vehicle 100 isdefined by a narrow cross section, as exemplified in vehicle bus 102.The bus 102 includes a first or top panel 110, a second or bottom panel112, and lateral sides 114 and 116. At the nose of the NEO vehicle 100is a leading edge 104, which is configured with a bevel to slope towardone or both the first or second panel 110, 112.

The example NEO vehicle 100 of FIG. 2 is shown in perspective view,illustrating a bevel 118 sloping from the leading edge 104 to the panel110. In some examples, the bevel 118 is angled at 20 degrees, andanother bevel opposite bevel 118 slopes toward panel 112. In someexamples, the angle is greater than or less than 20 degrees. Inexamples, the bevel 118 slopes at a first angle, whereas the oppositebevel slopes at a second angle different from the first angle.Furthermore, the bevel can slope at a constant angle on a flat surface,or can progress at a varying gradient toward the panels 110, 112. Othervariations on the surface of the bevel can also be employed, includingovoid-conical shape, pyramidal shape, etc., with the key feature beingthat the frontal area is sharply angled along the direction of travel.As described herein, the small cross section of the NEO vehicle 100, aswell as the sloping bevel from the leading edge 104, reduces drag on thevehicle 100 from atmospheric particles and aids in maintaining stableorientation in orbit.

FIG. 3 shows a cross-section of an example NEO vehicle 100 illustratingvarious components, including a radio frequency antenna 150 (e.g., aphased array). A computing platform 152 can include a processor, memorystorage, and/or various sensor types, such as attitude controlgyroscopes. A battery 154 or other storage system (e.g., capacitor,etc.) can be used to store power collected by solar panels in order to,for example, power the various components and the electronic engine 106of the NEO vehicle 100.

One or more optical imaging systems/lenses 156,158 are also included(e.g., variable field of view, multispectral imaging, etc.). The lenses156, 158 are configured to have a thickness sufficient to providedetailed imaging (e.g., a 1 m resolution at NEO altitudes) yet thinenough to fit within the vehicle bus 102, along with the various othercomponents. A folded light path contributes to reduced thickness of anoptical assembly, while a radar assembly can be made from an arraysimilar to the radio phased array antenna. Additionally oralternatively, the system can include a mechanical device to control theorientation of the lenses 156, 158 and or the antenna 150 to adjust thefocus of the respective system. A baffle 162 can be used to providestability as well as filtering stray light effects from non-imagedsources, supported by one or more posts 164. Each spacecraft isconfigured with sufficient area/volume to house one or more imagingsystems, such as two camera lenses 156, 158, and one or more baffles162. In some examples, a camera lens can be a 10 cm thick optical lenssystem, and a baffle external to the vehicle bus is used.

Many aspects of the spacecraft have equal applicability for systemsconfigured for image capture (e.g., optical data collection) and radarcapable spacecraft. In some examples, considerations related to size ofthe vehicle, weight, drag, power demands, as well as propellant needs,may change based on these and other factors. For example, in someembodiments, the cross-sectional area for an imaging satellite isgreater than that for a radar capable satellite (e.g., about 5 cm thickvehicle bus for radar satellite, compared with about 10 cm thick tohouse the camera optics).

Additional and alternative components may be included in the NEO vehicle100, such as radar or radio components, sensors, electronics bays forelectronics and control circuitry, cooling, navigation, attitudecontrol, and other componentry, depending on the conditions of theorbiting environment (e.g., air particle density), the particularapplication of the satellite (e.g., optical imaging, thermal imaging,radar imaging, other types of remote earth sensor data collection,telecommunications transceiver, scientific research etc.), for instance.In some examples, the system can include one or more passive and/oractive systems to manage thermal changes, due to operation of thecomponents themselves, in response to environmental conditions, etc. Thecomputing platform 152 can be configured to adjust the duty cycle of oneor more components, transfer power storage and/or use from a given setof batteries to another, or another suitable measure designed to limitoverheating within the NEO vehicle 100.

A fuel storage tank 160 is coupled with the engine 106 to generatethrust to counter the forces on the NEO vehicle 100 from drag, or toposition the vehicle in the proper orbit. The present and desired orbitcan be compared and any adjustments can be implemented by the computingplatform 152. For example, based on sensor data and/or a GlobalPositioning System (GPS) receiver, the computing platform 152 candetermine spatial information indicative of a current altitude of thesatellite, an orientation of the satellite relative to a terrestrialsurface, and a position of the satellite relative to other satellites.This data can be compared against a desired altitude, orientation orposition. If the computing platform 152 determines an adjustment isneeded, the ion engine 106 is controlled to generate thrust sufficientto achieve the desired altitude, orientation or position.

The ground stations must be numerous in order to achieve near continuousdownload of data. At 260 km and assuming that the vehicle antennas have90 degrees of steering, a vehicle can “see” a ground station within acircle on the earth of about 490 km and for about 60 seconds. Duringthis time, in order to avoid buffering camera data and at 1 meterresolutions, the downlink must transmit at a data rate of 400 Mbps aftererror coding is factored in. The inventors have determined that the mostfavorable link parameters to achieve low power consumption whilepreserving adequate link margins is to use the “Ku” band (e.g., about11.7-12.7 GHz) with a Quadrature Phase-Shift Keying (QPSK) modulationscheme has a symbol rate of 2 bits/symbol. In some examples, a BinaryPhase Shift Keying (BPSK) modulation scheme, and a symbol rate of 1bit/symbol. In examples, a Frequency-Shift Keying (FSK) modulationscheme can be employed.

In an example employing a QPSK modulation scheme operating in the Kuband, an 870 Mbps data rate can be achieved using 543.75 MHz ofbandwidth, consuming approximately 22.5 W of power. A suitable NEOvehicle antenna size is approximately 17.2×17.2 cm, as described withrespect to FIG. 4 . In such a transmission scheme, a modulation index ofapproximately 2 bits/symbol and using an error correction coding rate of0.75, a signal frequency at approximately 12.2 GHz (within the Ku band)with a transmission data rate of 870 Mbps achieves a 652.5 Mbps net datarate.

In an example using a BPSK modulation scheme operating in the Ku band, a875 Mbps data rate can be achieved using 1094 MHz of bandwidth,consuming approximately 3401 W of power with an antenna size ofapproximately 17.2×17.2 cm. A modulation index of approximately 1bit/symbol and using an error correction coding rate of 0.75, a signalfrequency at 12.2 GHz with a transmission data rate of 400 Mbps resultsin a 300 Mbps net data rate.

In accordance with the described transmission scheme, short revisittimes can be described above as “near-real time.” Traditional LEO andMEO satellites have revisit times of days to weeks, depending on thenumber of satellites in the constellation. Due to extremely highsatellite costs plus high launch costs, satellite constellations aretypically limited to a few to a few dozen satellites. Some proposedsystems include up to about 100 satellites, promising revisit times downto a day or so.

Near real-time revisit rates offer many advantages and solve manyproblems inherent in current satellite systems. One example is the“worst case” revisit time as compared to the average revisit time. Mostsatellites spend about half their orbit in earth's shadow (i.e., night)resulting in poor or useless images. Adding in cloud cover, up to 70% ofearth's surface, sand storms and perspective issues (e.g., images takenaround noon cast no shadow and are therefore more difficult tointerpret) reduce the number of useful images to about one fifth or lessof all images taken.

This sampling problem makes it difficult to plan image capture of acertain spot at a certain time. For many implementations, the averagetime to a useable image may not be as important as the worst case time,which we define as the time between images that meet a certain set ofcharacteristics (e.g., a specific location plus morning or evening, plusno cloud cover, etc.). In this example, getting images of a specificarea (e.g., a battlefield or a river flood plain) with a long revisittime constellation can make a worst case scenario push from days intoweeks. In this example, a system with a 3-day average revisit time couldbe overhead at night for several sequential passes, and then encountercloud cover or dust storms when it is finally overhead with correctlighting. So an average revisit time of 3 days can become a one ortwo-week worst case scenario, a delay that reduces or even eliminatesthe value of the images.

Conversely, with an exemplary revisit time of an hour, the current NEOsystem will generally have a vehicle overhead any spot on earth duringdaylight hours, many times every day. Furthermore, as clouds and duststorms are not stationary, the probability of having a NEO vehicle 100overhead during a break in the weather is further increased. Since thesestatistics are not a purely linear extrapolation of the average revisittimes (i.e., they are exponential), worst-case revisit times become muchmore manageable with the described low revisit time NEO system.

Images are only useful once they are conveyed back to systems on Earth.The NEO vehicle 100 includes a widespread array of receiving stationsrather than the normally low number of centralized receiving stationsfound in use with traditional satellite systems. For example, with threereceiving stations (e.g., US, Australia and Europe), a traditional LEOsatellite will be within transmission range approximately every 30minutes (90/3), at best. If the imagery data is available with aninherent delay of a week due to the long revisit time described above, afurther 30 minute delay is relatively small. In the event that asatellite is not in communication with a ground based station at thetime of imaging, the satellite imaging and processing components canbuffer the image data, and transmit to the next available ground station(e.g., when imaging an ocean or uninhabitable area).

However, for the current NEO satellite system, with average revisitrates of an hour or less, down to minutes or seconds in some examples,such a delay would be a large percentage of the goal. Therefore, datacan be downloaded from the NEO vehicles to a large network of low-costearth receiving stations to enable low-latency data downloads, ideallywith latency from time of taking to time of receiving on the order ofminutes to tens of minutes. In some examples, new image data is capturedat a very fast rate. To ensure proper downloading of the acquired data,the images should be transmitted at the rate they are created, orfaster. If the image data cannot be downloaded at a suitable rate, thendata and/or images may be buffered or discarded. Thus, to ensure acomplete image is downloaded and/or created, the imaging system shouldbe within range of a ground station, with a reasonable communicationslink, during an imaging event.

In one exemplary solution, receiving stations may be mounted atopcommercial cellular base stations, of which there are about 300,000 inthe US alone. Most such base stations are designed to support cellularcommunications radially outward. Therefore, an upwardly pointedradiation pattern can use the open area at the top of the base stationtower or on top of a suitable structure (e.g., a building, etc.),directing and receiving all energy to/from an orbiting NEO satellite andaway from any interference with the cellular signals.

In order to download sufficient data during an overpass of a single NEOsatellite and to meet the size, mass and cost targets of the NEOsatellite, a simple antenna with a relatively wide beam will enable arelatively large footprint on earth's surface. For example, a beam withfull width half max (FWHM) beam angle of 45° from 100 km altitude wouldhave a circular footprint about 200 km in diameter. Assuming thevehicle's orbital velocity is about 7.8 km/s, a useable receive time ofabout 26 seconds would result. A narrower beam would reduce this timewhile a wider beam would increase it.

Thus, in some examples, a phased array antenna is provided on both theNEO satellite and the ground based station. The beamwidths for theantennas are small (e.g., 6.4 and 3.2 degrees). This narrow beamwidthallows the antennas to achieve a high gain (e.g., 28.9 and 34.9 dB),which serves to reduce power consumption. Phased array technology hashistorically been used in military and space applications, but recentadvances in silicon technology are enabling highly integrated, costeffective solutions to be developed for next generation cell phonecommunications systems in commercial markets.

These disclosed narrow beamwidths, combined with speed and low altitude,allow both the ground station and spacecraft antennas to be steered asneeded to provide continual communications coverage (e.g., steered on afrequent or continual basis). In conventional systems, the gain of thephased array antenna is reduced as the scan angle increases, making itdifficult to scan very far with patch antennas. To overcome theshortcomings in the conventional systems, the antennas described hereinare designed with a scan angle of about +/−45 degrees. Additionally, theantenna gain goes up as the beam width narrows. The increased gain, andthus improved link budget and received signal level, can enable a higherorder of signal modulation, and thus a higher data transfer rate.

In order to ensure low-latency downloads, downloads may occur when avehicle is passing over long stretches of ocean or other “dead zones”,of which the oceans are the largest. In addition to ensuringavailability of sufficient receiving stations on islands, receivers mayalso be placed on ships or buoys to receive the images, which can thenbe transmitted to processing centers via traditional high capacitysatellites or fiber links.

In addition, NEO vehicles may include a vehicle-to-vehicle communicationsystem, such as with point-to-point laser systems. Using such aninter-vehicle link would enable very high-speed data rate transferbetween vehicles, enabling downloads to be handled by a vehicle otherthan the one collecting an image. Adding this flexibility to the systemhas several benefits, including filling dead-zone gaps, backupcapability if receivers are unavailable, and backup capability if adownlink transmitter on a NEO vehicle becomes disabled.

As the altitude increases, so does the power requirements for the radiosystem; however, the drag on the satellite decreases exponentiallyreducing the thrust requirements to overcome orbital decay, and thedestructive effects of high velocity impacts with atomic scaleatmospheric particles, such as Oxygen, decrease. In the presentinvention, it has been determined that an altitude of 260 km allows forabout 1 meter resolution of the earth's surface by optical imagers witha power requirement for the imager and radio combination that can besatisfied in a small, low drag profile vehicle that can be kept in orbitfor several years, as disclosed in the related application entitled“System For Producing Remote Sensing Data From Near Earth Orbit.” Uniqueto this system, the communications scheme only requires one groundstation every 420 km or so for a total of about 60 to 100 groundstations to cover the United States. Additionally or alternatively, upto about 850 ground stations can cover the Earth's land mass, and up toabout 1,100 ground stations can cover the Earth's surface (e.g.,including water borne or airborne antennas, etc.). At 260 km orbitalaltitude, the satellite will be able to “see” a ground station for about63 seconds before it must establish a new link with the next groundstation on its orbital path.

One component of the communications system is a planar antenna. FIG. 4shows an exemplary design for such an array. For example, the phasedarray antenna 150 includes a plurality of elements 151 designed tocommunicate with a ground based transponder. The elements 151 arearranged in a grid, with M number of elements 151 along they horizontalaxis times N number of elements 151 along the vertical axis to definethe phased antenna array. Although illustrated as an 8×8 element phasedarray antenna, any number of MxN elements 151 can be used to facilitatecommunications. For instance, a 16×16 element phase array antenna can beimplemented in a flat, square package in the NEO vehicle 100 (see, e.g.,FIG. 3 ). This configuration has the advantage of a low profileintegration into the satellite, as well as being steerable to “point”the downlink at the earth station improving link performance. Thebeamwidth is set for 6.4 degrees, and the array size is 17.1 cm×17.1 cm,small enough to fit within the “Cubesat” unit form factor of 2U (20cm)width on a vehicle having a 10 cm maximum vertical thickness (1 U) whilestill carrying the image package, engine, electronics, and solar panels.

To reduce the power requirements for the link, the ground stationutilizes a larger antenna that is also a phased array, in the presentinvention the array is 28×28 elements with an overall size of 34.5×34.5cm. This larger size has a higher gain than the antenna on the vehicle,which allows the satellite to transmit at a lower power. Like theantenna on the satellite, it is also steerable so that the maximumantenna gain can be “pointed” at the satellite.

In the present system, it is contemplated that the satellite will be “inview” of a ground station for about 60 seconds. During this window, theground station must establish a link with the satellite, download theimage data for at least a 60 second period of images, download telemetryfrom the satellite regarding operational status, and upload any commandsto the satellite. The ground station also forwards the downloaded dataand telemetry to a terrestrial wide area network for collection andprocessing at the satellite constellation control facility.

One feature of the present system is a design plan for locating groundbased stations in such a manner as to ensure continuous coverage. In theexample of FIG. 5 , plural ground based stations 240-240N can beseparated by a distance 124 to ensure the range 120-120N of adjacentstation antennas overlap 122. As shown, each ground based station240-240N has an approximate range 126, beyond which the associatedantenna is unable to receive data from a passing NEO vehicle 100. Thus,as each NEO vehicle 100 transmits information to a ground based station240-240N during flight, when the antenna array associated with groundbased station 240 reaches its limit, the range 120N from the antennaarray associated with adjacent ground based station 240N is configuredto “hand off” the information from the NEO vehicle 100. Such signaltransfer is described with respect to FIG. 6 .

FIG. 6 shows an example of satellites interacting with plural groundstations in accordance with aspects of this disclosure. As shown, aplurality of satellites 100A-100C are in a near earth orbit, asdescribed herein. A vehicle-to-vehicle laser communication system may beincluded to improve data download rates, flexibility and reliability.Each satellite 100A-100C is equipped with communications systems tocommunicate with other satellites (e.g., laser communications, radiocommunications, etc.). In the example of 90 satellites in an orbitalplane at about 1 minute intervals, distance between satellites will beapproximately 450 km. Since the horizon from 160 km altitude is morethan 1,000 km away, a laser communications system is capable ofproviding a direct link to multiple satellites in the same orbitalplane. Since the vehicles will be oriented along the orbital plane inorder to minimize drag, the control system for the inter-vehicle lasercommunications may be simple, for example, including possibly a fixedorientation.

For example, satellite 1008 can send and receive information tosatellite 100A via link 244 and with satellite 100C via link 246. In ahigh volume constellation with close spacing at low altitudes, line ofsight laser communications to neighbor vehicles will be effective. Inthe example of 90 satellites in an orbital plane at 1-minute intervals,distance between satellites will be approximately 450 km. Since thehorizon from 160 km altitude is more than 1,000 km away, a lasercommunications system is capable of providing a direct link to multiplesatellites in the same orbital plane with minimal atmospheric diffusioneffects at low power. Since the vehicles will be oriented along theorbital plane in order to minimize drag and their relative positionschange very slowly, the pointing system for the inter-vehicle lasercommunications may be relatively simple. Using such an inter-vehiclelink would enable very high-speed data rate transfer between vehicles,enabling downloads to be handled by a vehicle other than the onecollecting an image. Adding this flexibility to the system has severalbenefits, including filling dead-zone gaps, backup capability ifreceivers are unavailable, and backup capability if a downlinktransmitter on a NEO vehicle becomes disabled.

As shown in FIG. 6 , each satellite 100A-100C is configured to send andreceive information to and from ground based systems 240A-240C. Eachground based system 240A-240C is configured to communicate with anotherground based system via communication links 248, 250. For example,communications links 248 and 250 can be laser based, radio frequencytransmissions, wired connections, or a combination thereof. Thecommunication links may utilize dynamic beam shapes to maximize datadownload during each pass of satellites.

The system further includes a distributed earth receiver system relyingon a large number of receivers each downloading data during a satelliteoverpass. For instance, ground based systems 240A-240C are configured tocommunicate with satellites 100A-100C to send and receive informationvia communication links 242A-242C. Additionally or alternatively, aground based system can communicate with more than one satellite, orvice versa. As shown in FIG. 6 , ground based system 240A iscommunicating with satellite 100A via communications link 242A, and isalso configured to communicate with satellite 1008 via link 252. Inexamples, ground based station 240A can anticipate the arrival ofsatellite 1008 and adjust one or more antennas to facilitate datatransfer. The position of satellites within the orbit can be determinedbased on information stored in a database and available to each groundstation and/or satellite. The database can be updated in response todata received through earlier ground based station communications toimprove estimates of a given satellite's location, speed and/or otheroperational parameters. Moreover, communication between the ground basedstation 240A and satellites 100A and 1008 can occur simultaneously or insuccession.

Although the orbits are numerous, they are precisely known. Each groundstation is connected to a satellite ephemeris server that contains dataon when, where on the horizon, and on what trajectory the next satellitewill appear in view of any given ground station. Each ground stationthen, aims its antenna beam at the location in the sky where thesatellite is next expected to appear. In the same way, each satellitestores its own corresponding table of ground stations in order ofappearance, with instructions for where on the surface of the earth toaim its antenna to track the next ground station. In some examples, theinformation regarding location of the ground station is transmitted toeach satellite during a communication event. Additionally oralternatively, the location data includes information regarding thefrequency, power, Doppler shift, etc., associated with a particularsatellite. Access to such transmission characteristics further enhancesthe ability to track satellite movement, and allows a ground satelliteto anticipate the arrival of a satellite and prepare the antenna for aparticular transmission.

For example, both the ground station and the spacecraft can estimate theDoppler shift and compensate for it. Such anticipation and compensationfacilitates initial signal acquisition. For instance, if the receivingsystem knows to shift up by 100 kHz during signal acquisition, thesignal is acquired more efficiently and effectively. Moreover, theDoppler shift will change continuously as the spacecraft moves throughthe ground station coverage area, and the frequency shift will becontinuously tracked and corrected based on the spacecraft movement,speed, location, distance from the ground station, etc.

The electrically steerable nature of the antennas coupled with the factthat the location of both the ground station and satellite are knownallows for fast and efficient link establishment as the satellite movesover the earth. The ground station table on the satellite is updatedperiodically by the reverse link from the earth station. It is notedthat the uplink requirements are substantially less demanding than thedownlink since a very low data rate uplink, on the order of a fewthousand bits per second, is sufficient to keep the satellites onboardground station table up to date. In the event that the table fails, thesatellite reverts to an acquisition mode where the earth under theflight path is scanned for an acquisition signal from a ground station;when a signal is acquired, the satellite's tables are refreshed by theground station.

In some examples, a single ground station communicates with multiplesatellites. For instance, the satellites can transmit information in aunique frequency and/or with varying modulation schemes. To communicatewith satellites in different orbital positions, the ground station mayhave multiple antennas (e.g., steerable antenna) configured to track themovement of a satellite and/or the transmission characteristics of thesatellite.

In some examples, the ground station is portable. For instance, theground station can be mounted on a vehicle (e.g., a wheeled vehicle, anairborne vehicle, a watercraft, etc.) such that a communications linkcan be provided in an area that does not maintain a permanent receiver.Such a ground station can be deployed in disaster areas, conflict areas,and/or areas were imaging may be available for short periods. Portableground stations can communicate with other ground based stations viacellular data links and/or temporary communications cabling, as theenvironment and/or situation allows.

In the example of FIG. 7 , one or more NEO vehicles 100 can maintain anorbit 256 around the Earth 258, in accordance with the presentdisclosure. In one example, 90 satellites per necklace can be used,however more or fewer satellites per necklace may be appropriate for agiven application. For example, 45 satellites per necklace would spacethe vehicles at 2-minute intervals, while 180 would space vehicles at30-second intervals. As a person of ordinary skill will understand, theearth will rotate during the interval between arrivals of two sequentialNEOs, with that distance determined by the time separation between thesatellites. Different spacing distances may impact other subsystemdesigns such as optical imaging and radio links, but the concept remainsthat a NEO satellite system can provide relatively high rates ofcoverage. Although FIG. 6 shows three satellites in succession, anynumber of satellites into the tens of thousands can be employed in asatellite constellation, and can be aligned in a single direction oftravel in a single orbit, or may be traveling at angles with respect toeach other, and occupy multiple orbits (see, e.g., FIG. 7 ).

A communications system for use with a large constellation of imagingsatellites in near earth orbits has been described.

The system includes a large number of the low-cost, low-mass, lowaltitude NEO vehicles, thereby enabling revisit times substantiallyfaster than any previous satellite system. The NEO vehicles spendvirtually all of their orbit at the low altitude, high atmosphericdensity conditions that have heretofore been virtually impossible toconsider. Short revisit times at low altitudes enable near-real timeimaging at high resolution and low cost. The system further includes adistributed earth receiver system relying on a large number of receiverseach downloading data during a satellite overpass. The communicationlink may utilize optimized beam shapes to maximize data download duringeach pass. A vehicle-to-vehicle laser communication system may beincluded to improve data download rates, flexibility and reliability. Byoperating at such altitudes, the orbital mechanics have no impact on thespace junk issues of traditional LEO orbits and the system isself-cleaning in that any space junk or disabled craft will quicklyde-orbit.

1. (canceled)
 2. A plurality of ground stations configured to:communicate with one or more satellites; and perform a hand-off of theone or more satellites from a first ground station at a predeterminedfirst location to a second ground station at a predetermined secondlocation, wherein the one or more satellites occupies a region ofoverlap within a transmission range of the first ground station and atransmission range of a second ground station during the hand-off. 3.The plurality of ground stations of claim 2, further comprising: one ormore phased array antennae configured to transmit and receive data,wherein the one or more phased array antennae are electronicallysteerable to provide continuous coverage to the one or more satellites.4. The plurality of ground stations of claim 2, wherein the one or moresatellites operate at an orbit of 100 to 400 km.
 5. The plurality ofground stations of claim 2, further configured to perform a secondhand-off from the first or second ground station to a third groundstation of the plurality of ground stations.
 6. The plurality of groundstations of claim 5, wherein the one or more satellites occupy a secondregion of overlap within a transmission range of at least two of thefirst, second, or third ground stations during the hand-off.
 7. Theplurality of ground stations of claim 2, wherein communication with oneor more satellites is performed at a data rate of 300 Mbps or greater.8. The plurality of ground stations of claim 2, wherein the plurality ofground stations are arranged to ensure the transmission range of thefirst ground station overlaps with the transmission range of the secondground station.
 9. The plurality of ground stations of claim 2, whereinthe plurality of ground stations operate in a Ku frequency band.
 10. Theplurality of ground stations of claim 2, wherein at least one of thefirst ground station and the second ground station is co-located with acellular base station.
 11. The plurality of ground stations of claim 2,wherein the plurality of ground stations is further configured toreceive telemetry information from the one or more satellites and uploadinformation to the one or more satellites.
 12. The plurality of groundstations of claim 11, wherein the information uploaded to the one ormore satellites comprises hand-off information.
 13. The plurality ofground stations of claim 2, wherein the plurality of ground stations isfurther configured to communicate with a terrestrial wide area network.14. The plurality of ground stations of claim 2, wherein the pluralityof ground stations is further configured to communicate with a serverthat includes satellite data comprising satellite location andtrajectory information for the one or more satellites.
 15. The pluralityof ground stations of claim 2, wherein the plurality of ground stationsare separated by a distance to ensure adjacent station antennas have anoverlapping range.
 16. The plurality of ground stations of claim 2,wherein the plurality of ground stations are further configured toanticipate an arrival of the one or more satellites.
 17. The pluralityof ground stations of claim 16, wherein the plurality of ground stationsare further configured to adjust one or more antennae to facilitate datatransfer.
 18. A system comprising: one or more satellites; and aplurality of ground stations configured to: communicate with the one ormore satellites; and perform a hand-off of the one or more satellitesfrom a first ground station of the plurality of ground stations at apredetermined first location to a second ground station of the pluralityof ground stations at a predetermined second location, wherein the oneor more satellites occupy a region of overlap within a transmissionrange of the first ground station and a transmission range of a secondground station during the hand-off.
 19. The system of claim 18, furthercomprising a database including historical satellite location, speed, orother operational parameters.
 20. The system of claim 18, wherein theone or more satellites are configured to transmit data or images.
 21. Amethod for handing off a satellite between a first ground station and asecond ground station comprising: communicating with one or moresatellites; and performing a hand-off of the one or more satellites fromthe first ground station at a predetermined first location to the secondground station at a predetermined second location, wherein the one ormore satellites occupy a region of overlap within a transmission rangeof the first ground station and a transmission range of a second groundstation during the hand-off.